Missile device responsive to aerodynamic conditions

ABSTRACT

In combination with a ground-to-ground ordnance missile having a predictable trajectory during which the missile attains a supersonic velocity, improved means for obtaining an activating signal at a predetermined altitude prior to impact and after the missile has begun to decelerate, said means comprising in combination: first and second heat-responsive probes mounted to the exterior surface of said missile such that the aerodynamic heating of said probes during missile flight are highly independent of the surface conditions of the missile and missile geometry, each probe being located at the same distance aft of the missile nose so as to be exposed to the same aerodynamic heating conditions during missile flight, said probes being so constructed that said first probe has a heat capacity which is substantially different from the heat capacity of said second probe, said probes thereby having substantially different timetemperature relationships during missile flight, the heat capacities of said probes further being chosen in conjunction with predicted values of missile velocity, ambient air density and ambient air temperature throughout the missile trajectory so that the instantaneous temperature of said probes will become equal at said predetermined altitude, and probe-responsive means within said missile connected to said probes for producing an activating signal when the instantaneous temperatures of said probes become equal, said probe-responsive means including an electronic bridge-type circuit for sensing the temperatures of said probes by sensing their resistivities.

Elnited States Fatent [191 Pollin 11] 3,831,524 1451 Aug. 27, 1974MISSILE DEVICE RESPONSIVE TO AERODYNAMIC CONDITIONS [75] Inventor: IrvinPollin, Washington, DC.

[73] Assignee: The United States of America as represented by theSecretary of the Army, Washington, DC.

22 Filed: Feb. 21, 1957 21 Appl. No: 641,782

Primary Examiner-Benjamin A. Borchelt Assistant ExaminerC. T. JordanAttorney, Agent, or FirmEdward J. Kelly; Herbert Berl EXEMPLARY CLAIM Incombination with a ground-to-ground ordnance missile having apredictable trajectory during which the missile attains a supersonicvelocity, improved means for obtaining an activating signal at apredeter- FRQBE mined altitude prior to impact and after the missilehasbegun to decelerate, said means comprising in combination: first andsecond heat-responsive probes mounted to the exterior surface of saidmissile such that the aerodynamic heating of said probes during missileflight are highly independent of the surface conditions of the missileand missile geometry, each probe being located at the same distance aftof the missile nose so as to be exposed to the same aerodynamic heatingconditions during missile flight, said probes being so constructed thatsaid first probe has a heat capacity which is substantially differentfrom the heat capacity of said second probe, said probes thereby havingsubstantially different timetemperature relationships during missileflight, the heat capacities of said probes further being chosen inconjunction with predicted values of missile velocity, ambient airdensity and ambient air temperature throughout the missile trajectory sothat the instantaneous temperature of said probes will become equal atsaid predetermined altitude, and probe-responsive means within saidmissile connected to said probes for producing an activating signal whenthe instantaneous temperatures of said probes become equal, saidproberesponsive means including an electronic bridge-type circuit forsensing the temperatures of said probes by sensing their resistivities.

2 Claims, 3 Drawing Figures i RESPGNSE eases SAFETY a ALTITUDE a 0 FUZEARMIMG SPEED ,3 CHAMSai PUTER PATENTEDA BZ 3,881,524

ALTITUDE TEMPERATURE TIME PROBE RESPONSE MEANS J SAFETY 8a ALTITUDE &FUZE ARMING SPEED COM- /3 MECHANISM PUTER MEANS INVENTOR. lrvm PollinMISSILE DEVICE RESPONSIVE TO AERODYNAMICS CONDITIONS The inventiondescribed herein may be manufactured and used by or for the Governmentfor governmental purposes without the payment to me of any royaltythereon.

This invention relates to an application of aerodynamic heating ofordnance missiles, and more particularly to an application ofaerodynamic heating of ordnance missiles attaining supersonic speeds toperform useful functions therein.

An object of this invention is to apply aerodynamic heating to theperformance of useful functions.

Another object is to apply aerodynamic heating to the performance ofuseful functions for ordnance missiles.

Still another object is to apply aerodynamic heating to the performanceof useful functions for ordnance supersonic missiles of all types.

A principal object is the application of aerodynamic heating effects tothe safety and arming of supersonic missiles of all types.

An important object is the application of aerodynamic heating effects tothe safety and the arming of supersonic missiles of all types atpredetermined altitudes.

Another object is the employment of a continuous signal derived fromaerodynamic heating effects to accomplish safety and arming ofsupersonic missiles of all types.

A further object is to employ elements differing in their specific heatcapacities to initiate missile safety and arming at a predeterminedaltitude for a prescribed trajectory for supersonic missiles of alltypes.

A still further object is to employ elements differing in theirrespective heat capacities to initiate missile safety and arming at apredetermined altitude for a prescribed trajectory for a supersonicmissile independent of the missile acceleration or deceleration.

Another object of this invention is the application of aerodynamicheating effects to the fuzing of ordnance missiles.

Still another object is the application of aerodynamic heating effectsto the fuzing of ordnance supersonic missiles at predeterminedaltitudes.

A further object is the application of aerodynamic heating effects tothe fuzing at predetermined altitudes of ordnance supersonic missiles ofthe ground-toground, and airto-ground type.

A still further object is to employ elements differing in their specificheat capacities to initiate missile fuzing at a predetermined altitudefor a prescribed trajectory for an ordnance supersonic missile.

Still another object is to employ elements-differing in their respectiveheat capacities to initiate missile fuzing at a predetermined altitudefor a prescribed trajectory for a free-falling supersonic missile of theground-toground and air-to-ground type.

An important object of this invention is the application of aerodynamicheating, to the continuous measurement of instantaneous altitudes andspeeds of an ordnance missile.

Another object is the application of aerodynamic heating to thecontinuous measurement of instantaneous altitudes and speeds of anordnance missile at supersonic speed.

Still another object is the continuous measurement of instantaneousaltitudes and speeds for missile intelligence and telemetering.

A further object is to provide the constituent discriminatory elementsthat comprise the above objects singly or in any combination withemployment of a single pair or several pairs of temperature sensitiveprobes.

Briefly, this invention is based on the fact that missiles, attainingsupersonic speeds, will experience temperature changes during flight.Two temperaturesensitive probes having respectively different heatstorage capacities, when subject to the same aerodynamic heatingconditions, will yield distinctly different timetemperature curves. Loadmeans are coupled with the probes to interpret aboard the missile, or totelemeter, signals obtained from the probes, or to respond to apredetermined signal therefrom to accomplish a useful function at aselected altitude or velocity for a missile.

The specific nature of the invention as well as other objects, uses andadvantages thereof will clearly appear from the following descriptionand from the accompanying drawing, in which:

F [G l is a graphical representation of a free-fall, supersonic,ground-to-ground missiles altitude and the variations in temperature ofthe probes of the invention as a function of time for an exemplarytrajectory.

FIG. 2 is a schematic diagram of a system contemplated by the presentinvention.

FIG. 3 is a plan view of a portion of a missile incorporating a possiblemounting for one of the probes of this invention with certain partsbroken away.

Those familiar with ordnance missiles are aware of the aerodynamicheating effects on missiles during missile flight, especially thosemissiles attaining supersonic speeds. This invention applies theseheating effects to accomplish useful functions, and will be explained indetail with reference to a free-falling ground-to-ground ballisticmissile which attains supersonic speeds. lt will be understood, however,that the invention is not necessarily limited to such missiles.

It has been found that on descent free-falling supersonic missiles willexperience decreasing acceleration and ultimately deceleration becauseof an increase in atmospheric density amongst other factors. As a resultof this deceleration, the aerodynamic heating effects will be reduced sothat the missile may even start to cool. Two probes of different heatstorage capacities exposed to the same aerodynamic heating conditionswill yield distinctly different time-temperature curves. A probe with asmall heat capacity will have temperature values approximating those ofthe missile recovery temperatures and consequently will tend to cool atsubstantially the same time and at the same rate as the recoverytemperature. The instantaneous recovery temperature is the temperaturethat would be obtained by a thermally insulated missile if theinstantaneous aerodynamic heating conditions were continued over asufficient interval of time. Additional descriptions, considerations andmeasurements of recovery temperature may be found in National AdvisoryCommittee for Aeronautics, Technical Note 3623, March 1956. However, asecond probe having a higher heat capacity exposed to the sameaerodynamic heating will change in temperature less rapidly. Thus, atthe time the recovery temperature begins to decrease, the temperature ofthe probe with the smaller heat capacity is much higher than thetemperature of the other probe because of the latter's relatively slowresponse to aerodynamic heating, In fact, when the probe with thesmaller heat capacity begins to cool, the probe with the higher heatcapacity is still undergoing a temperature increase but will coolshortly thereafter.

Referring now to FIG. 1, a plot of a missiles trajectory from launch tomissile impact is presented by the solid line as a function of altitudeversus time. As will be apparent from the above taken in conjunctionwith FIG. 1, the two probes will heat during the major portion ofmissile trajectory and will cool during terminal portion of missileflight at respectively different rates. Line 11 shows the recoverytemperatures of a ground-to-ground missile as a function of time. Line12 denotes the temperatures attained by the lower heat capacity probeand line 14 temperatures attained by the higher heat capacity probe. Asshown and in practice the heat capacities of the probes may be selectedso that as the missile approaches the earth the temperatures of the twoprobes will become equal, as at T3, prior to missile impact.Furthermore, this equalization temperature T3 may selectively occur atany time during the aerodynamic cooling of the probes prior to missileimpact. Accordingly, by either mathematical computation or graphicalanalysis, the altitude at which probe temperature equalization occurscan be predetermined, selected, or preset for particular combinations ofprobes for a given missile trajectory. Calculations show that thealtitude at which the two probe temperatures become equal dependsprimarily on the val ues of the heat capacities of the probe. Similarly,the temperatures, the temperature differences, and rate of change oftemperatures for each probe at a selected altitude may be determined.The altitude at which probe temperature equalization occurs or theprobes assume definite temperatures, temperature differences or definiterates of change of temperature will be termed the fuzing altitude. Itwill be understood that, by means of suitable schemes for measuring andinterpreting the probe temperatures, various desired actions e.g.,arming or detonation can be automatically accomplished when the fuzingaltitude is reached.

The deceleration of some free-falling missiles and the accompanyingreduction in recovery temperatures begin at comparatively low altitudes.Therefore, for these missiles the probe temperatures must become equalshortly following the onset of the decrease in the recovery temperatureswhen temperature equalization determines the fuzing altitude. Hence,probes having relatively low heat capacities are required. The practicallower limit for the heat capacities corresponds primarily to thepractical lower limit for the probe thickness. However, it has beenfound, even at relatively low available fuzing altitudes, heatcapacities are obtainable that will result in probe cooling. Othervalues of heat capacities exist that will also provide a temperaturereduction even though equalization of the probe temperatures may not bepossible due to a limited amount of cooling that can take place in thetime intervals available. Moreover, in practice it is desirable that thetwo probes heat and cool at significantly different rates if possible,primarily because unless the two are cooling at substantially differentrates, considerations that may result in a small undesirable temperaturechange or error in either probe will cause a relatively largeundesirable change in the fuzing altitude. As the heat capacity of aprobe is increased, reduced rates of heating and cooling are obtained,and conversely, increased rates are obtained as the heat capacitydecreases, hence many combinations of heat capacities of the probes maybe utilized to arrive at probe temperature equalization at the samefuzing altitude. Thus, by decreasing the value of the heat capacity ofthe probe having the higher heat capacity, both the maximum temperatureattained and the temperature of this probe at the fuzing altitude willbe reduced. Similar results in fuzing altitude may be obtained byincreasing the value of the heat capacity of the probe having thesmaller heat capacity.

It is desirable to select a pair of probes for which the rate of changeof the difference between the two probe temperatures at the fuzingaltitude is approximately equal to the rate of change of the missilerecovery temperature. In this case, the difference between the heatcapacities is large and the fuzing altitude occurs almost immediatelyafter the maximum temperature is attained by the probe with the higherheat capacity. Similar considerations are taken into account when thefuzing altitude is determined by the respective temperatures of thestrips on their temperature differences or their respective rates ofchange of temperature for any particular altitude for a definitetrajectory as will be evident to those skilled in the art.

The probes 18 and 20 may be mounted so as to form part of the exteriorsurface or shell of the missile proper. The probes are located at thesame distance aft of the missile nose and thereby are exposed to thesame aerodynamic heating conditions. Probes l8 and 20 are preferablyidentical except for thickness and may be suitably mounted in recessedportions in the shell of missile l6, and as such are preferablyelectrically and thermally insulated from the shell and exterior ofmissile 16.

Referring now to FIG. 2, a free-falling supersonic ordnance missile isschematically shown at 16. The relatively high heat capacity probe 18and the relatively low heat capacity probe 20 may be positioned withrespect to the missile 16 substantially as shown for schematic purposes.However, the particular location of this probe on the shell of missile10 may be considered important. The aerodynamic heating of a missile isinfluenced by the type boundary layer. A turbulent boundary layer on themissile shell provides a greater aerodynamic heating than a laminarboundary layer. The boundary layer is always laminar in the immediateneighborhood of the missile nose and on proceeding towards the missiletail becomes turbulent. This transition of the boundary layer fromlaminar to turbulent flow is affected by such factors as the surfaceconditions of the missile, missile speed, and kinemetic viscosity of theambient atmosphere. Transition from a laminar to a turbulent boundarylayer moves toward the missile nose as the missile descends. The surfaceof the shell of missile 16 forward of the probes 18 and 20 may have aroughness at least comparable to a manufacturers painted article finishso that early transition from laminar to turbulent flow may arise atpositions toward the missile nose. Then the location of the probes mustbe such that the type of boundary layer in the vicinity of thefuzing-altitude is known in order to obtain high accuracy.

Temperature-sensitive probes having the form of a cylinder withhemispherical ends may also be used.

These probes are mounted so that the aerodynamic heating of the probeswould be highly independent from the surface conditions of the missileand its geometry. As illustrated in FIG. 3, (the direction of flightbeing in the direction of the enlarged arrow) relatively small bosses 22may extend from missile l6 and suitably receive probes 18 and 20. Thus,in this embodiment probes 18 and may assume a substantially cylindricalhemispherical configuration of suitable thickness and surface area toprovide the desired heat capacities. Suitable insulation should beprovided for the probes l8 and 20 from missile 16 and bosses 22 tominimize heat transfer between these parts. The heat ca pacities ofeither of the two types of probes l8 and 20 can be changed in the fieldwithout entering the missile interior by the use of a device that willserve to increase or decrease the effective thickness of the probes. Forexample, such a device may be a screw arrangement, the adjustment ofwhich is accessible from the missile exterior for adding or removinglamina or solely for insertion or withdrawal of screws of the samematerial as probes 18 and 20, thereby increasing or decreasing the heatstorage capacities of the probes. By these means, the variation of thefuzing altitude for a particular combination of probes over an intervalof ranges can be reduced to zero. Further, these means may serve toselectively alter the heat capacities of the probes to modify or changethe fuzing altitude for a specified trajectory or range.

For a given missile range, there is a family of trajectories which willobtain the desired range. Moreover, the trajectory characteristics for agiven missile range may be varied by changing the missiles burning time,cutoff angle, diving angle, etc. By suitable adjustment of theseparameters and others, the same fuzing altitude may be obtained for agiven interval of ranges. Similar considerations and factors affectfuzing altitude when the latter is determinate upon other probetemperatures and the respective rates of change of temperatures of theprobes.

Meteorologic conditions encountered by a missile during flight will havesome effect on fuzing altitude. An increase of the ambient density willtend to increase the air drag acting on the missile and thereby willreduce missile speed since the air drag is proportional to the productof the density and the square of the velocity. At high altitudes the airdrag acting on a free-falling missile is initially much less than thecomponent of the missile weight in the direction of the drag because ofthe small values of the ambient density. Consequently, the missileaccelerates on approaching ground until the air drag and the componentof the missile weight in the drag direction becomes equal. During thistime, the missile attains a very high speed and the rate at which thedensity increases becomes correspondingly large. For this reason, theair drag acting on a missile may become several times larger than themissiles weight. Under most situations the ambient density can bepredicted or calculated with a reasonable degree of certainty. However,if the actual density is not as predicted, the change in the fuzingaltitude due to a variation of the missile velocity accompanying achange of the ambient density has been found to be negligible for mostpractical applications.

For altitudes extending from sea level up to 26 miles, the ambienttemperature varies between about *70 and +90F. Although thetemperature-altitude relation varies with the geographical location andthe season, the temperature at altitudes below 100,000 feet may bepredicted with an accuracy of within i8F. This depends on theavailability of temperature-altitude data along the path taken by themissile, which is usually readily obtainable. If there is a deviationbetween the actual and predicted ambient temperatures, again thedeviation is usually harmless.

The effects of heat transmission by convection, conduction, andradiation on probes l8 and 20 may be made negligible by suitableinsulation, such as air or gas seals and by proper selection of surfaceconditions.

The high temperatures attained by the probes l8 and 20 limit thematerial from which they may be formed. Materials with low meltingpoints may not be used. Furthermore, the use of ferromagnetic materialis limited to temperatures below the Curie point since it then becomesparamagnetic and the specific heat capacity changes abruptly withtemperature. Since some alloys undergo a transition at a definitetemperature resulting in changes in the specific heat capacity, theiruse is limited. Materials that oxidize at high temperatures cannot beused. In addition, the probes 18 and 20 cannot be made of any materialfor which the thermal conductivity is undesirably low. Examples of thematerials that could be used are platinum and tantalum.

Probably the most simple and accurate method for determining the fuzingaltitude may be provided by comparing the electrical resistivities ofthe two probes l8 and 20 instead of directly comparing theirtemperatures. The electrical resistivity of a material is independent ofthe heat storage capacity and depends only usually linearly on itstemperature. Thus no calibration of thermometers or other specialtemperature sensitive or interpretive devices are required.

Load means, load circuitry or electronic sensing equipment may beemployed to function upon the desired signal from the probes l8 and 20and may comprise a conventional temperature response means or circuit 24utilizing each probe as an arm. It will be obvious to those skilled inthe art that other circuits or mechanisms can be used and should be usedfor the many applications of the present invention. The high order ofsensitivity of a bridge-type circuit may be desirable in practicebecause of the relatively small magnitude of the output signal derivedfrom the difference in temperatures of the probes l8 and 20. Theintelligence obtained by the probe temperature response means 24 maythen be applied to trigger or energize a safety and arming mechanism 26or any of a number of other mechanisms, such as a fuze 28 or altitudeand speed computer means 30.

To accomplish missile arming by the present invention the missilepreferably must have attained supersonic speeds during flight and notnecessarily during reentry. This requirement is substantially met by allmissiles of the supersonic variety which include those that areground-to-ground, air-to-ground, air-to-air, and ground-to-air. Asmentioned above safety and arming may be accomplished at that altitudewhen the probe temperatures become equal. This can be insured in theground-to-ground missile and the air-to-ground missile if missiledeceleration is present as in most freefalling type missiles. Safety andarming may also be adapted to occur at that altitude giving a definitecombination of probe temperatures which include the combination of thetemperature differences of the probes such as T for the relatively lowheat capacity probe and T for the relatively high heat capacity probe asillustrated in FIG. 1. This application of aerodynamic heating can beaccomplished in all types of ordnance missiles of the supersonic type byconventional temperature response means 24. Similarly, the rates ofchange of temperatures of the probes such as dIl/dt and dT2/dt (FIG. 1)at a particular altitude can be predicted and compared by conventionaltemperature rate of change response means or circuitry to accomplishmissile safety and arming for all type supersonic missiles. One of themost significant contributions of the present invention is the provisionof a continuous signal for safety and arming. This is achieved byconventional sensing equipment adapted to permit safety and arming onlyafter the respective probes experience a range of certain presettemperatures or preset rates of change in temperatures, or presetcombination of probe temperatures. Thus it will be obvious that armingcan take place at any point of a missile trajectory for all supersonicmissiles adapting the teachings of the present invention. The probetemperature rate of change response means may be substituted for theprobe temperature response means 24 or employed in combination with thelatter, as where the altitude and velocity determinations are employedin combination with safety and arming and fuzing, with suitableswitching means separating the two rendering one or the other or bothresponsive to the probes l8 and 20.

The essential requirements for practical and accurate fuzing of amissile that incorporates the present invention are that the missilepreferably must have been supersonic and have decelerated during reentryso that a temperature comparison or rate of change comparison of theprobes may be made during the terminal portion of flight immediatelyprior to missile impact. This is met by all free-falling ballisticground-to-ground missiles and most air-to-ground missiles whosetrajectory can be readily predicted. However, for a known or pre' dictedtrajectory all types of ordnance supersonic missiles can be effectivelyfuzed in a similar manner as missile safety and arming which wasexplained in detail above. In the light of the following disclosure,skilled persons will be able to construct a conventional computor 30 soas to obtain measurements of instantaneous missile speed and altitude.

The temperature variation AT of each probe during the time interval A!(see FIG. 1) is known to those skilled in the art and is where h is thelocal effective convective heat transfer coefficient in Btu/sec Fabsolute ft T,, is the recovery temperature F absolute;

y is the specific weight of probe, lb/ft C is the specific heat ofprobe, Btu/lb F;

1' is the probe thickness, ft;

T is the probe temperature, F absolute. Since instanh a(pv)"A and T =TKV where a, b, and K are constants, and p, v and T are the instantaneousvalues of ambient air density, missile speed, and ambient airtemperature at a selected alti tude, respectively. A is a function of Toand may be expressed as A (To).

Curves are predictable for p and To as functions of altitude y frommeterological data, and may be expressed t-aneous values of dT ldt, dT-/dt. T and T can be mea- .claims.

P=p(y) Substituting equations (8a) and (8b) into equation (7) will yieldmy) l o(y)l where the instantaneous values of h and T are measured bymeans of equation 1) and (2), as noted above. Equation (9) is now solvedby a reiterative process until the solution for y is found. According toequation (9) there is a unique value of y for every pair of h and T,,values. Substitution of To into equation (6) will yield theinstantaneous value of v. Thus, if h and T measurements are madecontinuously, continuous measurements may be made for y and v.Consequently, the values of y and v may be used to furnish missileintelligence or telemetered by any of the existing conventionaltechniques.

It will be apparent that the embodiments of the present invention may besuitably combined in many ways. For example, the safety and armingsystem 26 may be employed in conjunction with the fuzing system 28, andboth may be adapted to function along with the altitude and speedmeasuring system 30. Thus, from the basic contribution to the art by thepresent invention many ramifications are possible. Further, there areother controllable factors that may contribute to distinctly differentprobe time-temperature curves aside from probe heat capacities, such asprobe surface conditions and surface area exposed to aerodynamicfriction. These factors may also be employed to carry out the presentteachings of the present invention.

It will be apparent that the embodiments shown are only exemplary andthat various modifications can be made in construction and arrangementwithin the scope of the invention as defined in the appended 1 claim:

1. In combination with a ground-to-ground ordnance missile having apredictable trajectory during which the missile attains a supersonicvelocity, improved means for obtaining an activating signal at apredetermined altitude prior to impact and after the missile has begunto decelerate, said means comprising in combination: first and secondheat-responsive probes mounted to the exterior surface of said missilesuch that the aerodynamic heating of said probes during missile flightare highly independent of the surface conditions of the missile andmissile geometry, each probe being located at the same distance aft ofthe missile nose so as to be exposed to the same aerodynamic heatingconditions during missile flight, said probes being so constructed thatsaid first probe has a heat capacity which is substantially differentfrom the heat capacity of said second probe, said probes thereby havingsubstantially different time-temperature relationships during missileflight, the heat capacities of said probes further being chosen inconjunction with predicted values of missile velocity, ambient airdensity and ambient air temperature throughout the missile trajectory sothat the instantaneous temperatures of said probes will become equal atsaid predetermined altitude, and proberesponsive means within saidmissile connected to said probes for producing an activating signal whenthe instantaneous temperatures of said probes become equal, saidprobe-responsive means including an electronic bridge-type circuit forsensing the temperatures of said probes by sensing their resistivities.

2. The invention in accordance with claim 1, wherein each probe hassubstantially the form of a cylinder with hemispherical ends, the axisof said cylinder being parallel to the missile axis; and wherein saidmeans for mounting said probes comprises a pair of relatively smallbosses extending from the missile, each boss adapted to receive a probeat the probe end furthest from the missile nose, said probes and saidbosses being insulated from each other so as to minimize heat trans-

1. In combination with a ground-to-ground ordnance missile having apredictable trajectory during which the missile attains a supersonicvelocity, improved means for obtaining an activating signal at apredetermined altitude prior to impact and after the missile has begunto decelerate, said means comprising in combination: first and secondheat-responsive probes mounted to the exterior surface of said missilesuch that the aerodynamic heating of said probes during missile flightare highly independent of the surface conditions of the missile andmissile geometry, each probe being located at the same distance aft ofthe missile nose so as to be exposed to the same aerodynamic heatingconditions during missile flight, said probes being so constructed thatsaid first probe has a heat capacity which is substantially differentfrom the heat capacity of said second probe, said probes thereby havingsubstantially different timetemperature relationships during missileflight, the heat capacities of said probes further being chosen inconjunction with predicted values of missile velocity, ambient airdensity and ambient air temperature throughout the missile trajectory sothat the instantaneous temperatures of said probes will become equal atsaid predetermined altitude, and probe-responsive means within saidmissile connected to said probes for producing an activating signal whenthe instantaneous temperatures of said probes become equal, saidprobe-responsive means including an electronic bridge-type circuit forsensing the temperatures of said probes by sensing their resistivities.2. The invention in accordance with claim 1, wherein each probe hassubstantially the form of a cylinder with hemispherical ends, the axisof said cylinder being parallel to the missile axis; and wherein saidmeans for mounting said probes comprises a pair of relatively smallbosses extending from the missile, each boss adapted to receive a probeat the probe end furthest from the missile nose, said probes and saidbosses being insulated from each other so as to minimize heat transfertherebetween.